Jet engine for hypersonic intake velocities

ABSTRACT

A jet engine, which is suited for hypersonic intake velocities, has a rotating cascade in the intake duct for decelerating the hypersonic flow. The blade angles Beta 1, Beta 2 of the rotating cascade, the circumferential speed U of the rotor and the deceleration Delta w of the relative cascade flow are matched in accordance with cos Beta 2 (U . cos Beta 1)/(U Delta W . cos Beta 1), to ensure an axial direction of the absolute flow at the rotor cascade exit. A stationary vane cascade can be arranged downstream of the rotor cascade, in the case of high blade camber, for further deceleration and axial straightening of the flow. The intake rotor can be coupled to an engine accessory power unit or to a compressor arranged downstream of the intake. Controllable flaps can be provided to route the gas jet leaving the engine combustion chamber either to a turbine arranged downstream for driving the compressor or directly to the engine outlet nozzle. The intake rotor can also be coupled to an acceleration cascade arranged downstream of the combustion chamber; the acceleration cascade can also be preceded by a stationary vane cascade.

United States Patent Dettmering 1 March 6, 1973 [54] JET ENGINE FORHYPERSONIC FOREIGN PATENTS OR APPLICATIONS INTAKE VELOCITIES 611,44710/1948 Great Britain ..415/147 Inventor; Wilhelm Heinrich Dettmefing,am 696,007 8/1953 Great Britain ..4l5/l47 Chorusberg 26, Aachen, GermanyPrimary Examiner-Henry F. Raduazo [22] Fled: March 1970 Attorney-Craigand Antonelli l. N l9 4 8 [2]] App 0 1 57 ABSTRACT [30] ForeignApplication Priority Data A jet engine, which is suited for hypersonicintake velocities, has a rotating cascade in the intake duct for March14, 1969 Germany "P l9 l3 decelerating the hypersonic flow The bladeangles B B of the rotating cascade, the circumferential speed U of therotor and the deceleration Aw of the relative 60/3918 C cascade flow arematched in accordance with cos B [51] Int. Cl. ..F0ld 1/02, F02k 3/00 (Ucos BO/(U AW cos 3,), to ensure an axial [58] Field of Search ..60/270,226,269, 39.18 C; direction of the absolute flow at the rotor cascade415/181, 147, 143, 79 exit. A stationary vane cascade can be arrangeddownstream of the rotor cascade, in the case of high [56] Ref n e Cit dblade camber, for further deceleration and axial straightening of theflow. The intake rotor can be cou- UNITED STATES PATENTS pled to anengine accessory power unit or to a com- 2 693,675 11/1954 schafferw6069M; G pressor arranged downstream of the intake. Controlla- 2:702,9853/1955 Howell ..415/147 ble flaps can be Pmvided mute the gas letleaving 2,835,470 5/1958 Trowbridge et a1 ..60/39.18 c the enginembusti0n chamber either a turbine 2,939,017 5/1960 Teague, Jr. et a1...60/39.18 c ranged downstream for driving the compressor of 2,952,9739/1960 Hall et a1. ....60/39.18 C directly to the engine outlet nozzle.The intake rotor 2,945,672 7/1960 Wagner et al..... ....60/39.18C canalso be coupled to an acceleration cascade ar- 2,966,028 12/1960 Johnsonet al.... ..415/181 ranged downstream of the combustion chamber; the2,989,843 6/1961 Ferri ..60/226 acceleration cascade can also bepreceded by a sta- 3,422,625 1/1969 Harris tionary vane cascade2,947,139 8/1960 Hausrnann ..415/181 3,442,441 5/1969 Dettmering..415/181 14 Claims, 5 Drawing Figures (M t-c ("ml/ m ROTOR STATlONARYCASCADE VANE CASCADE c 11 7 I 11 u SHEET 1 or a PATENTED 5 I973 lNVEWILHELM DETT PAIENTEDHM: ems

SHEET 2 [IF 3 ON L) PATENTEDMAR 61% 3,719,428

SHEET 30F 3 ROTOR OUTLET STATIONARY VANE CASCADE INTAKE M l L ROTOR JETENGINE FOR HYPERSONIC INTAKE VELOCITIES BACKGROUND OF THE INVENTION Thepresent invention relates to a jet engine, and more particularly, to ajet engine suitable for hypersonic intake velocities.

To ensure satisfactory compressor and combustion chamber operation injet engines with hypersonic intake velocity, it has been necessary, upto now, to decelerate the hypersonic flow in the air air intakes tosubsonic velocity. With conventional engines, this has led tocompression shocks resulting in static pressure losses and thus to areduction of total engine efficiency.

The maximum losses occur in a normal compression shock and will increasewith the Mach number upstream of the shock. For this reason,deceleration is effected by one or more compression shocks acting in anoblique direction and causing considerably lower static pressure lossesor by means of a continuous and thus shock loss-free decelerationachieved with the aid of a wedge-shaped or conoidal bullet. This method,however, necessitates flow-tuming, thereby leading to an increase of theexternal drag of the engine and thus to net thrust losses. Therefore,the above-described deceleration can only be partially realized andwill, at higher flight speeds, e.g., approximately Mach 2.5, not besufficient to avoid considerable static pressure losses due to anexcessive Mach number immediately upstream of the occurring normalcompressionshock.

A further deceleration, which causes relatively slight static pressurelosses due to flow reversal and simultaneous area restriction, islimited by the set-on of supersonic flow rule, according to which ahypersonic flow in the intake duct can develop only if the mass flowpenetrating a normal compression shock in the intake duct does notexceed the capacity of the downstream throat area, otherwise the totaldeceleration downto subsonic velocity is effected in a normal shockupstream of the duct, resulting in correspondinglyhigh losses. Thepermissible area restriction and, thus also, the maximum possibledeceleration of the supersonic flow becomes higher as the intakevelocity C, (Mach number M becomes higher.

SUMMARY OF THE INVENTION It is an aim of the present invention toovercome the problems and disadvantages that have been encountered inthe prior art. y

it is a further aim of the present invention to provide 8 maximumreduction of the hypersonic velocity of engine intake flow, combinedwith minimum losses.

The underlying problems are solved in accordance with the presentinvention by utilizing the relationship that the minimum-lossdeceleration of a hypersonic flow which is made possible by arearestriction becomes higher as the Mach number of this flow immediatelyupstream of the deceleration becomes greater.

The foregoing problems are solved in accordance with the. presentinvention by arranging a rotating cascade in the intake duct fordecelerating the hypersonic flow. Due to the circumferential speed U ofthis rotating cascade, the relativeflow velocity W, in the cascadechannel is increased, since the circumferential velocity U is added, byvector addition, to the absolute intake velocity C which is normal tothe cascade plane. Thus, a Mach number M of the cascade flow is obtainedand is higher than the Mach number M, of the absolute intake flow. Asstated above, it will thus be possible to achieve a higher deceleration,A W=W, -W with a corresponding change in Mach number M=M in the cascadechannel while maintaining the supersonic speed, i.e., without a normalcompression shock, which will result in a corresponding static pressureincrease.

If, by means of suitable blade camber angles, simultaneously and inaddition a flow turning against the direction of the circumferentialvelocity is effected, there will result a considerably reduced absoluteexit velocity C, with Mach number M after vector substraction of thecircumferential velocity at the cascade channel exit. With an intakevelocity of W, having a Mach number M, 2, an arrangement of the intakerotor in accordance with the present invention resulted in a 25 percentincreased deceleration as compared to the conventional arrangementwithout a rotor cascade. Correspondingly, the arrangement in accordancewith the present invention resulted in the static pressure ratio in thedownstream normal compression shock being increased by 12 percent.

In accordance with an embodiment of the present invention, the bladeangles [3 3,, the circumferential speed U of the rotor and thedeceleration AW of the relative cascade flow are matched in accordancewith the function which will ensure an axial direction of the absoluteflow at the rotor cascade exit. This feature prevents any exchange ofenergy with the rotor so that the latter will not need any drivingenergy with the exception of the energy to overcome friction.

A further feature of the present invention increases the flow turningeffect in the rotor to such a degree, by means of a high blade camber,as to impart a circumferential component to the rotor exit flow. In thisconnection, a stationary vane cascade is arranged downstream of therotor cascade for further deceleration and axial straightening of theflow. Due to the more intensive turning of the supersonic flow in therotor cascade, in addition to the pressure increase by area restrictionof the cascade channels as abovedescribed, flow energy is also extractedin the form of shaft power, since an impulse is induced in the rotor inthis case. Through this measure, the possible shockfree deceleration AWW, W of the hypersonic flow is still further increased. With an absoluteflow inlet Mach number of M 3, an increase in the deceleration ofapproximately 32 percent is obtained as compared to the version withouta rotor cascade. In addition, a corresponding increase of the staticpressure ratio in the downstream normal shock of approximately 40%results.

In accordance with a further embodiment of the present invention, theintake rotor can be coupled to an engine accessory power unit for thetransfer of power thereto. As a result of this arrangement, the energyextracted from the flow is indirectly returned to the jet engine.

If no accessory drives are to be powered, the intake rotor can becoupled. to a compressor arranged downstream of the intake and therebytransfer the energy extracted from the intake flow to it in accordancewith a still further embodiment of the present invention. By includingthe compressor into the energy transformation taking place in theintake, an excessive compression of the working medium is avoided.Viewed as a system, the cascade rotating in the supersonic range and thecompressor coupled to it receive no external energy, thus the totalenthalpy between the inlet and exit of the system remains constant. Atthe high flight Mach numbers under consideration, the pressure increase,which is achieved solely by the lowloss in flow deceleration in theintake rotor-compressor system, is already so high that an additionalcompression effected by the power transferred from the turbine mightlead to an efficiency reduction of the engine, as the temperature riseassociated with it will reduce the possible heating margin of thecombustion chamber with a given constant maximum temperature.

Another embodiment in accordance with the present invention includescontrollable flaps which can be actuated so as to route the gas jetleaving the engine combustion chamber either to a turbine arrangeddownstream of the combustion chamber fOr driving the compressor ordirectly to the engine outlet nozzle; in this way, the flow energyextracted by the turbine is controlled, since only very low or even nopower is required from the turbine if the compressor is driven by theintake rotor.

In a further development in accordance with the basic idea of thepresent invention, the energy extracted from the flow within thehypersonic zone at the intake is returned to the engine only downstreamof the combustion chamber. Therefore, in an embodiment encompassing thisdevelopment the intake rotor is coupled to a rotating accelerationcascade arranged downstream of the combustion chamber. The flowacceleration in this cascade is achieved by pressure reduction due to anarea increase as well as by work transfer due to intensive turning bythe relative flow.

According to still another feature of the present invention, theacceleration cascade downstream of the combustion chamber can bepreceded by a vane cascade. The circumferential component of theabsolute velocity produced in this cascade suffices to effect the returnto an axial direction of the absolute velocity at the rotor outlet. Thisarrangement is particularly advantageous because it permits control ofperformance exchange by vane adjustment.

, Still in accordance with the present invention, a vane cascade can bearranged downstream of the rotor cascade located aft of the combustionchamber. This vane cascade serves to straighten the flow from the rotorcascade, which flow has a circumferential component, into an axialdirection.

As soon as the gas leaving the combustion chamber is heated to such adegree that it is at least partially ionized, it can also be acceleratedby electrical or magnetic fields. The present invention also proposes toproduce an electrical or electromagnetic field in the flow ductfof theengine downstream of the combustion chamber, whereby an ionized flowmedium is accelerated. The energy required for building up such anacceleration field is produced by a generator driven by the intakerotor.

To permit the use of an engine in accordance with the present inventionfor missile propulsion in a speed range as wide as possible, thepossibility of an adaptation to the respective speed ranges duringflight would be desirable. The foregoing is achieved by means of andadjustable rotor and stator blades of the deceleration stage arranged inthe intake duct. This relatively simple measure leads to a considerableextension of the speed range in which the missile can operate withadvantageous efficiency.

According to a further proposal, the rotor and stator blades of theacceleration cascade arranged downstream of the combustion chamber aremade adjustable, thereby similarly extending the speed range in whichfavorable efficiency can be achieved. In a still further embodiment inaccordance with the present invention, at least one coupling can beprovided by which the intake rotor is connected to or disconnected fromother rotating parts such as, for example, the accessory drive,compressor or acceleration cascade during operation.

In a further embodiment, adjustable flaps are ar ranged in the intakeduct for routing the intake flow upon selection either to the intakerotor or through a duct bypassing the intake rotor. This feature isespecially. useful in cases where the engine is also intended forsubsonic operation.

BRIEF DESCRIPTION OF THE DRAWING These and further aims, features andobjects of the present invention will become more apparent from thefollowing description when taken in conjunction with the accompanyingdrawing which shows, for purposes of illustration only, severalembodiments in accordance with the present invention and wherein:

FIG. 1 is a schematic view of an intake of a jetengine in accordancewith the present invention with an intake rotor and the associated bladescheme in addition to the velocity triangle diagrams at the cascadeinlet and outlet;

FIG. 2 is a schematic arrangement similar to FIG. 1 but with a vanecascade downstream of the intake rotor in addition to the correspondingvelocity triangle diagrams;

FIG. 3 is a schematic arrangement of the front section of a jet enginein accordance with the present invention and shows the intake rotorcoupled to the first stage of a downstream compressor along with theentropyenthalpy diagram for this jet engine;

FIG. 4 is a schematic view of another embodiment of the jet engine inaccordance with the present invention wherein the intake rotor iscoupled to an acceleration cascade arranged downstream of the combustionchamber, along with the corresponding entropy-enthalpy diagram; and

FIG. 5 is a schematic view of a still further embodiment in accordancewith the present invention, wherein the intake rotor, the compressor andan acceleration cascade arranged downstream of the combustion chamberare connected via a shaft, and adjustable flow control flaps areprovided to adapt the engine to the operating condition.

Referring now to the drawings wherein like numerals are used todesignate like parts throughout all the views and, in particular to FIG.1, there is shown one embodiment of the flow intake of a jet engine inaccordance with the present invention essentially consisting of aconoidal member 1 arranged at the center of an outer casing 2. To reducethe intake flow losses, a tip 1a of the conoidal member 1 protrudes in aconventional manner substantially beyond front end 2a of the outercasing 2. At the front end 20, an annular flow channel 3 is providedwith a rotor cascade 4 which rotates about center line 5. The frontlinesof oblique shocks developing at the intake are designated by the numeral6.

In FIG. 1, there is also shown the blade scheme,of rotor 4 along withthe corresponding velocity triangle diagrams upstream and downstream ofthe rotor 4, wherein:

C absolute velocity U circumferential speed of the rotor W relative flowvelocity in the cascade channel and wherein the subscripts used are:

condition of undisturbed flow in front of the conoidal member I=condition upstream of the rotor cascade II condition downstream of therotor cascade The arrangement shown in FIG. 2 is very similar to thatshown in FIG. 1 and described above. However, due to the higher camberof the rotor blade airfoils and the resulting flow velocity downstreamof the rotor 4, which has a circumferential component, a vane cascade 7is arranged downstream. As a result of the camber of vanes 7, the flowvelocity C, downstream of the vane cascade is straightened in an axialdirection. Auxiliary power means 10 are operatively connected to theintake rotor 4.

The jet engine shown in FIG. 3 has an intake in accordance with the oneshown in FIG. 2. Intake rotor 4 is rigidly connected to the first stage8 of an axial compressor by means of shaft 9. Additional rotors of theaxial compressor are designated by the numeral 11, while numeral 12designates a combustion chamber downstream of the compressor. Thecircled Roman numerals I, II, III, IV represent the various stages ofthe jet engine and correspond to the circled Roman numerals in theaccompanying entropy-enthalpy diagram.

FIG. 4 shows still a further embodiment of the jet engine in accordancewith the present invention which is intended for extremely high speedranges, e.g., Mach 2.5. Intake rotor 4 is mounted on shaft 9a which can,in turn, be coupled to shaft 9b by means of clutch 13. The clutch 13 canbe engaged and disengagedduring operation. Shaft 9b carries anacceleration rotor 14 which is located downstream of the combustionchamber 12. The vanes 7 are arranged downstream of the intake rotor 4.

The acceleration cascade 14 is located downstream of the combustionchamber 12 and is preceded by outlet vane 15 in which a circumferentialcomponent is thereby induced into the axially directed flow, The circledRoman numerals are used to designate the conditions of the jet engine asfollows:

O Condition of undisturbed flow in front of the conoidal member I=Condition upstream of the intake rotor II Condition downstream of theintake rotor IIl= Condition downstream of the intake vane IV= Conditiondownstream of the combustion chamber or, alternatively, conditionupstream of the outlet vane V Condition upstream of the accelerationvane VI Condition downstream of the acceleration vane VII Flow conditionafter leaving the jet engine The above conditions have been accordinglyentered in the accompanying entropy-enthalpy diagram in FIG. 4.

FIG. 5 shows another embodiment of the jet engine in accordance with thepresent invention wherein an intake rotor 4, compressor 8, and a turbine16 which is arranged downstream of the combustion chamber 12 are mountedon common shaft 9. To permit the jet engine to be adapted to variousinflow velocity ranges, swivelling flaps 17 are provided in the intakeduct which, upon selection, are used to direct the intake air either tothe intake rotor 4 or, through a duct, to bypass the intake rotor.Similar flaps 18 are arranged in the outlet duct for routing the floweither to the turbine 16 or to bypass it.

The movement of flaps 17 and 18 is carried out in a manner such that,with low intake velocities, the intake rotor 4 is covered by flaps l7and thus the flow is routed through the bypass duct, while flaps 18 openthe passage to the turbine 16. At high intake velocities, the pattern isreversed, that is the flow is routed through the intake rotor 4 and theturbine 16 is bypassed. Between these two extreme positions, any desiredintermediate position can be selected.

While I have shown and described several embodiments in accordance withthe present invention, it is to be understood that the same issusceptible of numerous changes and modifications as will be known to aperson skilled in the art, and I, therefore, do not wish to be limitedto the details shown and described herein but intend to cover all suchchanges and modifications as are encompassed by the scope of the presentinvention.

lclaim:

1. An air intake arrangement for hypersonic air flow duct means foraccepting hypersonic air flow entering a jet engine; said arrangementcomprising: an air duct means for accepting hypersonic flow, a freelyrotatable rotor means adjacent the forward end of said duct means, ashaft supporting said rotor means about a longitudinally extending axis,said rotor means including circumferential spaced blading about saidrotor means, said blading extending radially outwardly from adjacent theaxis across said duct means, the adjacent blades of said bladingdefining a plurality of constricting flow paths inclined to the axis ofrotation in a downstream direction whereby the air flow is constrictedas the air flows downstream along paths rearwardly from between the mostforward sections to between the most rearward sections of adjacent blademeans, said air flow constriction causing a reduction in the axialvelocity of air flow as the hypersonic flow decelerates along the flowpaths downstream, the flow velocity in a direction parallel to said flowpaths being greater immediately downstream of the most forward sectionsof said blade means than the flow velocity in the axial directionimmediately upstream of said most forward sections, and means forcontrolling the rotational velocity of said rotor means, saidlast-mentioned means comprising the geometrical configuration of theblade means whereby static pressure losses are minimized during thedeceleration by the constriction of the hypersonic flow through therotor means.

2. An arrangement according to claim 1, wherein the inlet angle (3 andthe exit angle 9,) of the blades of .the rotor means with respect to aplane extending perpendicular to the rotor axis, the circumferentialspeed (U) and the deceleration (AW) of the flow through the blades ofthe rotor means are matched in accordance with the relationship:

COS B2= (U cos ,8,)/(U AW cos (3,)

3. An arrangement according to claim 1, wherein stationary vane cascademeans are located downstream of the rotating rotor means for furtherdecelerating the flow and for straightening the flow into asubstantially axial direction when the flow leaving the rotating rotormeans has a substantially circumferential component.

4. An arrangement according to claim 1, wherein the shaft of the rotormeans is operatively connected to the engine for driving an auxiliarypower means.

5. An arrangement according to claim 4, wherein the inlet angle (3,) andthe exit angle (69 f the blades of the rotor meanS with respect to aplane extending perpendicular to the rotor axis, the circumferentialspeed (U) and the deceleration (AW) of the flow through the blades ofthe rotor means are matched in accordance with the relationship:

cos B2= cos BO/(U AW cos 3,

6. An arrangement according to claim 4, wherein stationary vane cascademeans are located downstream of the rotating rotor means for furtherdecelerating the flow and for straightening the flow into asubstantially axial direction when the flow leaving the rotating rotormeans has a substantially circumferential component.

7. An arrangement according to claim 1 wherein the shaft of the rotormeans is connected to a compressor operatively arranged in the enginefor supplying power to said compressor.

8. An arrangement according to claim 7, wherein the inlet angle (3,) andthe exit angle (B of the blades of the rotor means with respect to aplane extending perpendicular to the rotor axis, the circumferentialspeed (U) and the deceleration (AW) of the flow through the blades ofthe rotor means are matched in accordance with the relationship:

9. An arrangement according to claim 7, wherein stationary vane cascademeans are located downstream of the rotating rotor means for furtherdecelerating the flow and for straightening the flow into asubstantially axial direction when the flow leaving the rotating rotormeans has a substantially circumferential component.

10. An arrangement according to claim 7, wherein the shaft of the rotormeans is operatively connected to the engine for driving auxiliary powermeans.

11. An arrangement according to claim 2, wherein the blades areadjustable.

12. An arrangement according to claim 3, wherein the vane cascade meansincludes adjustable blades.

13. An arrangement according to claim 1, wherein the velocity of the airleaving the rotor means is supersonic.

14. An arrangement according to claim 1, further comprising auxiliaryload power take off means attached to said shaft.

1. An air intake arrangement for hypersonic air flow duct means foraccepting hypersonic air flow entering a jet engine; said arrangementcomprising: an air duct means for accepting hypersonic flow, a freelyrotatable rotor means adjacent the forward end of said duct means, ashaft supporting said rotor means about a longitudinally extending axis,said rotor means including circumferential spaced blading about saidrotor means, said blading extending radially outwardly from adjacent theaxis across said duct means, the adjacent blades of said bladingdefining a plurality of constricting flow paths inclined to the axis ofrotation in a downstream direction whereby the air flow is constrictedas the air flows downstream along paths rearwardly from between the mostforward sections to between the most rearward sections of adjacent blademeans, said air flow constriction causing a reduction in the axialvelocity of air flow as the hypersonic flow decelerates along the flowpaths downstream, the flow velocity in a direction parallel to said flowpaths being greater immediately downstream of the most forward sectionsof said blade means than the flow velocity in the axial directionimmediately upstream of said most forward sections, and means forcontrolling the rotational velocity of said rotor means, saidlast-mentioned means comprising the geometrical configuration of theblade means whereby static pressure losses are minimized during thedeceleration by the constriction of the hypersonic flow through therotor means.
 1. An air intake arrangement for hypersonic air flow ductmeans for accepting hypersonic air flow entering a jet engine; saidarrangement comprising: an air duct means for accepting hypersonic flow,a freely rotatable rotor means adjacent the forward end of said ductmeans, a shaft supporting said rotor means about a longitudinallyextending axis, said rotor means including circumferential spacedblading about said rotor means, said blading extending radiallyoutwardly from adjacent the axis across said duct means, the adjacentblades of said blading defining a plurality of constricting flow pathsinclined to the axis of rotation in a downstream direction whereby theair flow is constricted as the air flows downstream along pathsrearwardly from between the most forward sections to between the mostrearward sections of adjacent blade means, said air flow constrictioncausing a reduction in the axial velocity of air flow as the hypersonicflow decelerates along the flow paths downstream, the flow velocity in adirection parallel to said flow paths being greater immediatelydownstream of the most forward sections of said blade means than theflow velocity in the axial direction immediately upstream of said mostforward sections, and means for controlling the rotational velocity ofsaid rotor means, said last-mentioned means comprising the geometricalconfiguration of the blade means whereby static pressure losses areminimized during the deceleration by the constriction of the hypersonicflow through the rotor means.
 2. An arrangement according to claim 1,wherein the inlet angle ( Beta 1) and the exit angle ( Beta 2) of theblades of the rotor means with respect to a plane extendingperpendicular to the rotor axis, the circumferential speed (U) and thedeceleration ( Delta W) of the flow through the blades of the rotormeans are matched in accordance with the relationship: cos Beta 2 (U .cos Beta 1)/(U - Delta W . cos Beta 1)
 3. An arrangement according toclaim 1, wherein stationary vane cascade means are located downstream ofthe rotating rotor means for further decelerating the flow and forstraightening the flow into a substantially axial direction when theflow leaving the rotating rotor means has a substantiallycircumferential component.
 4. An arrangement according to claim 1,wherein the shaft of the rotor means is operatively connected to theengine for driving an auxiliary power means.
 5. An arrangement accordingto claim 4, wherein the inlet angle ( Beta 1) and the exit angle ( Beta2) of the blades of the rotor meanS with respect to a plane extendingperpendicular to the rotor axis, the circumferential speed (U) and thedeceleration ( Delta W) of the flow through the blades of the rotormeans are matched in accordance with the relationship: cos Beta 2 (U .cos Beta 1)/(U - Delta W . cos Beta 1)
 6. An arrangement according toclaim 4, wherein stationary vane cascade means are located downstream ofthe rotating rotor means for further decelerating the flow and forstraightening the flow into a substantially axial direction when theflow leaving the rotating rotor means has a substantiallycircumferential component.
 7. An arrangement according to claim 1,wherein the shaft of the rotor means is connected to a compressoroperatively arranged in the engine for supplying powEr to saidcompressor.
 8. An arrangement according to claim 7, wherein the inletangle ( Beta 1) and the exit angle ( Beta 2) of the blades of the rotormeans with respect to a plane extending perpendicular to the rotor axis,the circumferential speed (U) and the deceleration ( Delta W) of theflow through the blades of the rotor means are matched in accordancewith the relationship: cos Beta 2 (U . cos Beta 1)/(U - Delta W . cosBeta 1)
 9. An arrangement according to claim 7, wherein stationary vanecascade means are located downstream of the rotating rotor means forfurther decelerating the flow and for straightening the flow into asubstantially axial direction when the flow leaving the rotating rotormeans has a substantially circumferential component.
 10. An arrangementaccording to claim 7, wherein the shaft of the rotor means isoperatively connected to the engine for driving auxiliary power means.11. An arrangement according to claim 2, wherein the blades areadjustable.
 12. An arrangement according to claim 3, wherein the vanecascade means includes adjustable blades.
 13. An arrangement accordingto claim 1, wherein the velocity of the air leaving the rotor means issupersonic.